1. Field Of The Invention
This invention relates to a three-axis control angular momentum stabilized spacecraft, and to the re-orientation of such a spacecraft from the transfer orbit attitude to the on-station attitude.
2. Prior Art
Unmanned spacecraft are typically launched from the earth by means of three distinct types of launch vehicles. A first type, utilizing an unmanned launch rocket, utilizes a technique of direct injection whereby the launch vehicle propels the spacecraft directly into its final on-station orbit. In the case of a geosynchronous orbit located approximately 22,300 miles above the earth, direct injection techniques require substantially large launch vehicles, even for moderate size spacecraft.
The ability to place a spacecraft into a geostationary orbit is important because when on-station at that point, the spacecraft appears to remain fixed relative to a point on the surface of the earth. Hence, geostationary orbits are important for communications spacecraft and weather spacecraft which can be accurately positioned relative to ground-based antennas or provide continuous coverage of the same portion of the earth for scientific studies.
Direct injection capability has not been used because of the increased cost of launch vehicles and mission uncertainties. Hence, an alternative technique of placing the spacecraft into an initial orbit has been utilized with the spacecraft then undergoing a transfer orbit coupled with an appropriate firing of the spacecraft apogee motor to boost it into a synchronous orbit. The use of the transfer orbit launch sequence allows the use of smaller launch vehicles, such as the Atlas-Centaur or, in the 1980 time frame, the space shuttle.
In the transfer orbit, various attitude maneuvers are carried out to position the spacecraft for apogee motor firing used to boost it into a synchronous orbit as well as maintaining telemetry capabilities with the earth.
The transfer orbit maneuvers can be categorized in two basic groups, depending on the type of stabilization which the spacecraft will ultimately use in a synchronous orbit. If the spacecraft is to be spin stabilized in its on-orbit orientation, then precession maneuvers to position the spacecraft for apogee motor firing are conventionally carried out by first spinning up the spacecraft in the transfer orbit and allowing that spin rate to be maintained throughout the precession maneuvers; once complete, the spacecraft can be precessed to an earth coverage attitude.
The prior art is replete with techniques for carrying out such precession maneuvers on spin-stabilized spacecraft as typified by U.S. Pat. No. 3,294,344, to Rosen, et al., and U.S. Pat. No. 3,758,051, to Williams. Those two patents disclose a series of precession and velocity control maneuvers. The spin axis of the spacecraft is precessed to achieve a new attitude distinct from the initial attitude for re-orientation during the transfer orbit.
A second class of spacecraft are those which are controlled about the three principal axes while in an in-orbit mode. Hence, the spacecraft of this class are not in a spinning configuration while on-station but use concepts of stored bias momentum generally along the pitch axis together with thrustor firings to maintain a three-axis stabilized orientation.
Conventionally, in the prior art, such spacecraft are launched into a transfer orbit by a booster, such as an Atlas-Centaur or the space shuttle in combination with a spin-stabilized upper stage, after being spun up to achieve spin stabilization in the transfer orbit. The spacecraft remains spin stabilized in the transfer orbit with precession maneuvers carried out to re-orientate the spacecraft prior to apogee motor firing. After final orbit is achieved, solar arrays and communications antennas are deployed. The prior art is replete with various techniques for effectuating transfer orbit maneuvers of such bias momentum spacecraft. The actual use of spacecraft, such as the Communications Technology Satellite (CTS), presents a well-documented example of a bias momentum, three-axis stabilized spacecraft which has undergone precession maneuvers. This spacecraft has been reported in the literature, for example, in AIAA Paper No. 72-5800, "A High-Powered Communications Technology Satellite for 12/14 GHz Band," by Franklin and Davidson.
In such spacecraft, the technique of re-orientation is accomplished on the spacecraft, which is initially spinning about its axis of maximum moment of inertia in a spin-stabilized mode. When the precession maneuver is to take place, the spacecraft is despun to remove all angular momentum about the initial spin axis. Sun acquisition is accomplished in two steps: The spacecraft is first controlled with thrustors to capture the roll axis to the sun, and then a 90.degree. rotation is performed to align the yaw axis with the sun. The momentum wheel is then spun up while the yaw axis is kept pointing to the sun with thrustors. This technique requires a coordinated precession maneuver utilizing a programmed search maneuver to effect the desired orientation. Hence, the system requires an array of sensors, gyros and thrustors, all coordinated to maintain the spacecraft in a proper orientation during the despin portion, the search portion and when angular momentum is slowly imparted back into the system.
A second technique which has been proposed is to completely despin the spacecraft by means of thrustors with the momentum wheel remaining in a de-energized state. Under the control of various sensors, such as earth horizon sensors and sun sensors, the spacecraft is torqued by thrustors such that the roll axis is pointed toward the sun. In this orientation, the momentum wheel is then spun up, and when the sun, spacecraft and earth are in a co-linear orientation, the spacecraft is slowly rotated about the pitch axis until the yaw axis points toward the earth. The spacecraft is then processed until the pitch axis is normal to the orbit plane. This technique, while eliminating the coordinated momentum maneuver of the CTS system, requires hardware which will not be required for any other phase of the mission. For example, this technique requires that sophisticated rate gyros be utilized during the orientation maneuvers together with additional sensors for the orientation maneuver. This is a severe weight penalty for a one-time sequence of operations.
A third technique has been proposed in U.S. Pat. No. 3,940,096, to Keigler, et al. This technique, like the prior concepts, is applicable to a bias momentum stabilized spacecraft. The technique is used to convert a spin-stable configuration about the maximum moment of inertia axis to three-axis stabilization by a 90.degree. reorientation such that the direction of the thrust vector axis of the apogee motor and that of the axis of the momentum wheel will be interchanged while despinning the spacecraft. As in all of the concepts for three-axis spacecraft, the starting point is a spacecraft in a spin-stabilized configuration which is to be precessed into an orientation for final stabilization and deployment of solar arrays, antennas and the like. The re-orientation maneuver as set forth in the '096 patent consists of the steps of having the spacecraft first spin stabilized about its axis of maximum moment of inertia so that the spin axis is in alignment with the angular momentum vector of the vehicle. This is shown schematically in that patent as spin axis 18 for angular momentum vector H. The spacecraft is then despun to a second spin rate such that at the second spin rate, the angular momentum of the vehicle is substantially equal to the angular momentum stored by the momentum wheel when in an in-orbit condition. During both of these two steps, the momentum wheel is de-energized and the angular momentum vector of the spacecraft is generally pointing, as shown in FIG. 1, along the orbit normal.
the third step in the '096 process is to spin up the momentum wheel at a rate to provide for a reduction in the angular momentum of the body as the axis of the momentum wheel becomes parallel to the total system angular momentum vector, which remains fixed in space. Residual damping of the spacecraft results in only a small pitch rate existing as the final momentum in the body following the maneuver.
The disadvantage of this system is that it is not applicable to a spacecraft whih is not spinning about its axis of maximum moment of inertia. Contemporary spacecraft configurations having large deployable antenna farms and solar arrays will spin about an axis of minimum inertia because of constraints in packaging imposed by the launch vehicle configuration. Essentially, the vehicle shroud which covers the spacecraft during the launch sequence has a maximum size for aerodynamic and vehicle moment of inertia limitations. Once launched and the shroud is ejected, the spacecraft will spin about a principal axis during the transfer orbit to provide gyroscopic stabilization and ensure sufficient solar power for telemetry and command. However, the transition problem remains because initially the spacecraft will be spinning about an axis which is perpendicular to the momentum wheel axis.